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Abstract:

Disclosed is an aircraft, configured to have a wide range of flight
speeds, consuming low levels of power for an extended period of time,
while supporting a communications platform with an unobstructed
downward-looking view. The aircraft includes an extendable slat at the
leading edge of the wing, and a reflexed trailing edge. The aircraft
comprises a flying wing extending laterally between two ends and a center
point. The wing is swept and has a relatively constant chord. The
aircraft also includes a power module configured to provide power via a
fuel cell. The fuel cell stores liquid hydrogen as fuel, but uses gaseous
hydrogen in the fuel cell. A fuel tank heater is used to control the
boil-rate of the fuel in the fuel tank. The fuel cell compresses ambient
air for an oxidizer, and operates with the fuel and oxidizer at pressures
below one atmosphere.

Claims:

1. An aircraft, comprising: a hydrogen source; an oxygen source; a
control system configured to actively control and vary the pressures at
which the hydrogen and oxygen are supplied for reaction in the fuel cell,
wherein the control system is configured to actively control and vary the
reaction pressures of the fuel cell based on the altitude of the
aircraft; one or more aircraft engines configured to provide all
necessary thrust for the aircraft to fly; and a fuel cell configured to
react hydrogen from the hydrogen source with oxygen from the oxygen
source to generate power, wherein the fuel cell is configured to react
the hydrogen with the oxygen, the hydrogen and oxygen each being supplied
at a reaction pressure of less than one atmosphere absolute; wherein
power generated by the fuel cell powers the one or more aircraft engines;
and wherein the fuel cell is further configured to provide sufficient
power for the one or more aircraft engines to enable the aircraft to
loiter at altitude with the hydrogen and the oxygen being supplied for
reaction in the fuel cell at reaction pressures of less than one
atmosphere absolute.

2. The aircraft of claim 1, wherein the hydrogen is supplied at a
reaction pressure of no more than 11 psia.

3. The aircraft of claim 2, wherein the hydrogen is supplied at a
reaction pressure of no more than 10 psia.

4. The aircraft of claim 2, wherein the oxygen is supplied at a reaction
pressure of no more than 6 psia.

5. The aircraft of claim 2, wherein the increment of the fuel reaction
pressure from the oxidizer reaction pressure is no more than 5 psi.

6. The aircraft of claim 5, wherein the increment of the fuel reaction
pressure from the oxidizer reaction pressure is no more than 4 psi.

7. The aircraft of claim 1, wherein the oxygen is supplied at a reaction
pressure of no more than 6 psia.

8. The aircraft of claim 7, wherein the increment of the fuel reaction
pressure from the oxidizer reaction pressure is no more than 5 psi.

9. The aircraft of claim 8, wherein the increment of the fuel reaction
pressure from the oxidizer reaction pressure is no more than 4 psi.

10. The aircraft of claim 1, wherein the aircraft is configured for
continuous operation in conditions equivalent to an altitude of
55,000-70,000 feet.

11. The aircraft of claim 1, and further comprising solar cells
configured to provide power to the aircraft when the sun illuminates the
solar cells.

12. The aircraft of claim 1, wherein the control system is configured to
actively control and vary the reaction pressures of the fuel cell based
on the power requirements of the aircraft.

Description:

[0001] The present application is a continuation of application Ser. No.
11/973,091, filed Oct. 5, 2007, which is a divisional of application Ser.
No. 10/600,258, filed Jun. 20, 2003, now U.S. Pat. No. 7,281,681, issued
Oct. 16, 2007, which is a continuation-in-part of application Ser. No.
10/073,828, filed Feb. 11, 2002, now abandoned, which is a divisional of
application Ser. No. 09/826,424, filed Apr. 3, 2001, now U.S. Pat. No.
6,550,717, issued Apr. 22, 2003, which claims priority from two U.S.
provisional patent applications, Ser. No. 60/194,137, filed Apr. 3, 2000,
and Ser. No. 60/241,713, filed Oct. 18, 2000, each of which are
incorporated herein by reference for all purposes.

BACKGROUND OF THE INVENTION

[0002] This invention relates generally to aircraft and their component
systems, and, more particularly, to improved high-performance aircraft
systems capable of high-altitude stationkeeping within tight altitude and
perimeter boundaries for extended periods of time.

[0003] A worldwide expansion in the demand for communication bandwidth is
driving up the bandwidth requirements between satellites and
ground-stations. One way to increase this satellite-to-ground bandwidth
is to interpose one or more high-altitude platforms (HAPs) configured for
relaying signals between the two. A HAP allows for lower power
transmissions, narrower beamwidths, as well as a variety of other
advantages that provide for greater bandwidth. However, due to a
demanding set of design requirements, years of design efforts at creating
highly effective HAPs are only now beginning to come to fruition.

[0004] In particular, it is desirable to have a stratospheric aircraft,
capable of carrying a significant communications payload (e.g., a payload
of more than 100 kg that consumes more than 1 kw of electric power), that
can remain aloft for days, weeks or even months at a time. This flight
capability will preferably be maintainable even in zero or minimum
sunlight conditions where solar power sources have little functionality.
Also, the aircraft is preferably remotely pilotable to limit its weight
and maximize its flight duration.

[0005] The communications payload preferably is configured to view
downward over a wide, preferably unobstructed field of view. The aircraft
will preferably be capable of relatively high-speed flight that is
adequate to travel between its station and remote sites for takeoff
and/or landing to take advantage of benign weather conditions. At the
same time, the aircraft preferably is capable of maintaining a tight,
high-altitude station in both high-wind and calm conditions, thus
requiring relatively high-speed and relatively low-speed flight, and a
small turning radius while maintaining the payload's downward-looking
(and preferably upward-looking for some embodiments) view. To meet these
stringent design specifications, the performance of the aircraft's power
system, flight control system and airframe configuration and are all
preferably improved over prior practice.

Power Systems

[0006] Conventional aircraft are typically powered using aviation fuel,
which is a petroleum-based fossil fuel. The prior art mentions the
potential use of liquid hydrogen as a fuel for manned airliners and
supersonic stratospheric flight. There is also 25-year-old prior art
mentioning the possibility of using liquid hydrogen as fuel for a
stratospheric blimp.

[0007] U.S. Pat. No. 5,810,284 (the '284 patent), which is incorporated
herein by reference for all purposes, discloses an unmanned,
solar-powered aircraft that significantly advanced the art in
long-duration, stratospheric aircraft. It flies under solar power during
the day, and stores up additional solar power in a regenerative fuel cell
battery for use during the night to maintain its station. The fuel cell
battery is a closed system containing the gaseous elements of hydrogen
and oxygen that are dissociated from, and combined into, water.

[0008] The aircraft disclosed in the '284 patent is an unswept,
span-loaded, flying wing having low weight and an extremely high aspect
ratio. Multiple electric engines are spread along the wing, which is
sectionalized to minimize torsion loads carried between the sections.
Most or all of the sections contain a hollow spar that is used to contain
the elements used by the fuel cell. Large fins extend downward from inner
ends of the sections. The wings contain two-sided solar panels within
transparent upper and lower surfaces to take maximum advantage of both
direct and reflected light.

[0009] The above-described technologies cannot provide for long-duration,
high-altitude flights with tight stationkeeping when the available solar
power is highly limited.

Flight-Control Components

[0010] Various components are known for use in controlling flight. Each
component has unique advantageous and disadvantageous characteristics.

[0011] Many present-day small aircraft and some sailplanes use simple
flaps to increase camber and obtain higher lift coefficients, and hence,
adequate lift at lower speeds. Such flaps are typically retracted or
faired to reduce drag during high-speed flight, and also during
turbulence to reduce the maximum G loads that the wing will then
experience. An important characteristic of the use of flaps, or of the
use of highly cambered airfoils designed for high lift, is that the
extended flap or highly cambered airfoil provides the wing with a large
negative pitching moment. This affects both overall vehicle stability and
the wing's torsional twisting. Indeed, for high aspect-ratio wings, the
twist at the wing's outer portions due to a negative pitching moment can
pose severe structural and flight control problems.

[0012] Airliners use both leading edge slats and sophisticated flaps, such
as slotted or Fowler flaps, to widen their speed range. Small planes
employ slats that open automatically when needed. Hang gliders have
employed flexible airfoil tightening to decrease camber for high-speed
flight. Some work has been done with flexible flap material that unrolls
and pulls back from the rear of the wing. Some aircraft feature wings
characterized by a sweep that can be varied in flight, even turning the
entire wing so that it is not perpendicular to the flight direction
during high-speed flight.

[0013] For maintaining low-speed flight without stalling, large solid or
porous surfaces that hingedly swing up from a wing top in low-speed
flight to potentially stabilize vortices immediately behind them, are
known. This might provide an increased lift coefficient before stall is
reached. Various vortex generators and fences are used to delay the onset
of a stall or to isolate the portion of a wing that is stalled.
Furthermore, various stall warning/actuators allow aircraft to operate
relatively close to their stall speed. Additionally, some combinations of
airfoils and wing configurations feature gentle stalls and so the vehicle
can be operated at the stall edge without abrupt drag increases or lift
decreases during the onset of a stall. Experimental aircraft have even
employed rotary devices to permit low-speed flight, with mechanisms that
restrict rotary moment and decrease drag or potentially augment lift when
at higher speeds the wing provides the main lift. Many of the above
mechanisms provide this increased low-speed control at the expense of
weight and reliability.

[0014] In some high-tech aircraft, highly-active control is used to
maintain stable operation over a wide range of speeds and orientations.
This emulates the flying characteristics of natural fliers that change
wing and airfoil geometry. In aircraft, such systems are complex,
potentially heavy, and expensive, as well as fault-intolerant.

[0016] Furthermore, there is an inherent relationship between an
aircraft's overall airframe geometry and the design of its airfoils and
control surfaces. Typical aircraft offset negative (i.e., nose-down)
pitching moments through the use of tail moments (i.e., vertical forces
generated on the empennage with a moment arm being the distance from the
wing to the empennage) or through the use of a canard in front of the
wing that, for pitch stability, operates at a higher lift coefficient
than the wing and stalls earlier. Tails mounted in the up-flow of wingtip
vortices can be much smaller than tails positioned in the wing downwash,
but there are structural difficulties in positioning a tail in the
up-flow.

[0017] Commercial airliners address the high coefficient of lift (CL)
requirements for landing and takeoff with a complex array of slats and
flaps that are retracted during high-speed flight to lower drag and
gust-load severity. A rigid wing structure, and pitch controllability
from the tail's area and moment arm, permit this approach. However, this
approach is contrary to the requirement that the present aircraft carry
fuel adequate to last for extended periods of time, and still be
economical.

[0018] The very special requirements and technological challenges for the
aircraft of the present invention have not been met by existing aircraft
designs. Accordingly, there has existed a definite need for a lightweight
aircraft capable of both stationkeeping and flight over a wide range of
speeds, that consumes low levels of power for an extended period of time,
that supports a communications platform with a wide, unobstructed view,
and that is characterized by simplicity and reliability. Embodiments of
such an invention can serve as high altitude platforms. Embodiments of
the present invention satisfy various combinations of these and other
needs, and provide further related advantages.

SUMMARY OF THE INVENTION

[0019] The present invention provides aircraft, aircraft components and
aircraft subsystems, as well as related methods. Various embodiments of
the invention can provide flight over a wide range of speeds, consuming
low levels of power for an extended period of time, and thereby
supporting a communications platform with an unobstructed
downward-looking view, while and having simplicity and reliability.

[0020] In one variation, a wing of the invention is characterized by
having adequate camber to achieve a lift coefficient of approximately 1.5
at the Reynolds number experienced by sailplanes or flexible-winged
stratospheric aircraft. The wing defines a leading edge and a trailing
edge, and the trailing edge includes either a reflexed portion or a
trailing edge flap that can extend upward. Either the reflexed portion or
the flap is configured to provide the wing with a pitching moment greater
than or equal to zero in spite of the camber. This feature advantageously
allows for low-speed flight with a flexible wing in many embodiments.

[0021] This feature is augmented by an extendable slat at the leading edge
of the wing. These features, in combination, provide for an excellent
coefficient of lift of the wing, typically increasing it by more than
0.3, and preferably by 0.4 or more, at airspeeds just above the stall
speed. Using its retractability, the slat can become part of the wing's
airfoil that is otherwise defined by the wing's camber. Slats are
convenient because they have a negligible or beneficial effect on a
wing's pitching moment. While flaps might help increase the CL more
than slats, they do so at the cost of a big increase in negative pitching
moment that potentially, requires heavy, drag-producing countermeasures
for compensation.

[0022] In another variation of the invention, an aircraft comprises a
flying wing extending laterally between two ends and a center point,
substantially without a fuselage or an empennage. The wing is swept and
has a relatively constant chord. The aircraft also includes a power
module configured to provide power for the aircraft, and a support
structure including a plurality of supports, where the supports form a
tetrahedron. This tetrahedron has corners in supportive contact with the
wing at structurally stiff or reinforced points laterally intermediate
the center point and each end. The tetrahedron also has a corner in
supportive contact with the wing's center point, which is also
structurally stiff or reinforced. Advantageously, the flying wing is
configured with a highly cambered airfoil and with reflex at a trailing
edge. The wing is also configured with slats. These features provide many
embodiments with the capability of high-altitude flight with a wide range
of speeds.

[0023] A third variation of the invention is an aircraft, and its related
power system, for generating power from a reactant such as hydrogen. The
power system includes a fuel cell configured to generate power using a
gaseous form of the reactant, the fuel cell being configured to operate
at a power-generation rate requiring the gaseous reactant to be supplied
at an operating-rate of flux. The power system also includes a tank
configured for containing a liquid form of the reactant, wherein the tank
includes a heat source for increasing a boiling-rate of the reactant. The
tank is configured to supply its reactant to the fuel cell at a rate
determined by the boiling-rate of the reactant, and the heat source is
configured to increase the boiling rate of the reactant to a level
adequate for supplying the resulting gaseous reactant to the fuel cell at
the operating-rate of flux. An advantage of such an aircraft is that it
provides for a minimized system weight, volume and complexity, while not
excessively sacrificing power generation.

[0024] In a fourth variation of the invention, the power system of the
third variation includes a tank that comprises an inner aluminum tank
liner having an outer carbon layer, an outer aluminum tank liner having
an outer carbon layer, and connectors extending between the inner and
outer aluminum tank liners to maintain the aluminum tank liners' relative
positions with respect to each other. The volume between the inner and
outer tank liners is evacuated to minimize heat transfer between the
contents of the tank and the outside environment. The connectors between
the inner and outer layers are configured with holes in their walls to
minimize direct heat-conduction between the contents of the tank and the
outside environment.

[0025] In a fifth variation of the invention, an aircraft includes a
hydrogen source, an oxygen source and a fuel cell configured to combine
hydrogen from the hydrogen source and oxygen from the oxygen source to
generate power. The fuel cell is preferably configured to combine the
hydrogen and the oxygen at less than one atmosphere of pressure, and more
preferably at roughly 2-3 psia. This advantageously allows stratospheric
flight with simpler fuel cell technology.

[0026] Preferred embodiments of the above aspects of the invention, and
various combinations of their features, provide for unmanned aircraft
capable of flying in the stratosphere, in a stationkeeping mode, carrying
a payload of more than 100 kg that consumes more than 1 kw of electric
power, and remaining aloft for a significant period of time while being
able to operate from a remote site where takeoff/landing weather is
benign.

[0027] Other features and advantages of the invention will become apparent
from the following detailed description of the preferred embodiments,
taken in conjunction with the accompanying drawings, which illustrate, by
way of example, the principles of the invention.

BRIEF DESCRIPTION OF THE DRAWINGS

[0028] FIG. 1A is a perspective view of a first embodiment of an aircraft
embodying features of the present invention, the aircraft having a
cowling removed to expose a fuel tank that the cowling conceals.

[0029] FIG. 1B is a front elevational view of the embodiment depicted in
FIG. 1A, having its cowling in place.

[0030] FIG. 1C is a right side elevational view of the embodiment depicted
in FIG. 1B.

[0031] FIG. 1D is a top plan view of the embodiment depicted in FIG. 1B,
rotated by 90 degrees.

[0032]FIG. 2 is a system diagram of a fuel cell system for the embodiment
depicted in FIG. 1A.

[0033]FIG. 3 is a partial cross-sectional view of the fuel tank's wall in
the embodiment depicted in FIG. 1A.

[0034]FIG. 4 is a partial cross-sectional view of a cross cell connector
used in the fuel tank's wall depicted in FIG. 3.

[0035]FIG. 5 is a cross-sectional view of a wing on the embodiment
depicted in FIG. 1A.

[0036]FIG. 6 is a cross-sectional view of a wing on a variation of the
embodiment depicted in FIG. 1A.

[0037]FIG. 7A is a top plan view of a third embodiment of an aircraft
embodying features of the present invention.

[0038]FIG. 7B is a rear elevational view of the embodiment depicted in
FIG. 7A.

[0039]FIG. 8 is a top plan view of a fourth embodiment of an aircraft
embodying features of the present invention.

[0040] FIG. 9 is a top plan view of a fifth embodiment of an aircraft
embodying features of the present invention.

[0041] FIG. 10A is a top plan view of a sixth embodiment of an aircraft
embodying features of the present invention.

[0042] FIG. 10B is a rear, elevational view of the embodiment depicted in
FIG. 10A.

[0043] FIG. 11A is a bottom plan view of a seventh embodiment of an
aircraft embodying features of the present invention.

[0044] FIG. 11B is a front, elevational view of the embodiment depicted in
FIG. 11A.

[0045] FIG. 11C is a bottom plan view of a variation of the embodiment
depicted in FIG. 11A.

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS

First Preferred Aircraft Embodiment

[0046] A first preferred, high-performance aircraft embodiment 101,
capable of high-altitude stationkeeping within tight altitude and
perimeter boundaries for extended periods of time, according to the
present invention, is shown in FIGS. 1-4. The aircraft includes a wing
103, an empennage 105 and a plurality of motors 107. The empennage is
preferably suspended from the wing on an extension 109 to provide the
moment arm necessary to control pitch and yaw. Thus, the extension's
length will be based on the empennage's surface area and the needed
pitching and yawing moments.

[0047] A fuel tank 111 is suspended below the wing using trusses and/or
wires. A payload section 113 containing a communications payload 115
extends forward from a lower portion 117 of the fuel tank, and is
suspended using trusses, wires and/or supports 118. Preferably, the
aircraft includes a cowling or fuselage portion 119 (not shown in FIG. 1
to expose contents) that forms a single aerodynamic body enclosing the
fuel tank and payload section.

[0048] Preferably, the wing 103 is unswept, extending 200 feet tip-to-tip.
The wing preferably has a constant 10 foot chord, and thus an aspect
ratio of 20. The wing thus has an aspect ratio on the order of 20. Port
and starboard sides of the wing are each equipped with an inboard portion
121 having no dihedral and an outboard portion 123 having a positive
dihedral. The wing is torsionally flexible to limit the overall aircraft
weight.

First Preferred Aircraft Embodiment

Fuel and Power Systems

[0049] Each side of an inboard portion 121 of the wing mounts four
electric motors 107, and each side of an outboard portion 123 of the wing
mounts five electric motors, for a total of 18 electric motors. With
reference to FIG. 2, preferably the aircraft is powered by a hydrogen-air
fuel cell system that uses gaseous hydrogen as fuel. The system includes
a fuel cell 131 that combines a reactant of gaseous hydrogen with oxygen
and outputs electric power and water. The fuel cell powers an inverter
133 that runs a motor 135 that drives a compressor 137 to compress
outside air to provide oxygen for the fuel cell. The air and hydrogen
combine in the fuel cell to create the power both for the compressor's
inverter, and for an inverter 139 to run a propellor motor.

[0050] The fuel cell of the preferred embodiment can be configured to
operate with a fuel of gaseous hydrogen at reaction pressures (i.e., the
pressure of the hydrogen gas when it reacts with the oxygen via a
membrane electrode assembly) above 1 atmosphere, such as approximately 15
psi or higher. However, unlike typical hydrogen-powered systems, which
are designed with complex thermal and mechanical systems to operate at
reaction pressures of greater than one atmosphere, the present embodiment
is preferably designed to operate at reaction pressures ranging below 1
atmosphere, and possibly down to even 2 or 3 psia, when the aircraft is
at cruise altitudes. This feature can significantly reduce the cost and
weight of the power generation system while increasing its reliability
during high-altitude flight.

[0051] The hydrogen reacted by the fuel cell is preferably derived from a
fuel source comprising liquid hydrogen that is stored in the fuel tank
111. Storing the fuel as a liquid provides for the fuel to be stored in a
volume that is small enough to fit reasonable aircraft shapes.
Preferably, the cryogenic container(s) necessary to carry the fuel are
relatively lightweight. Other known hydrogen sources such as gaseous
hydrogen tanks are within the scope of the invention.

[0052] As noted above, the fuel cell reacts air as an oxidizer, the air
being compressed from ambient air. Preferably the air is taken into the
aircraft at any given cruise flight condition so that storage and
retrieval are not necessary. This allows for continuous operation (i.e.,
operation for an unlimited period of time within the fuel capacity of the
aircraft) of the aircraft at stratospheric flight conditions.

[0053] The oxidizer source for the aircraft preferably comprises an inlet
for ambient air and preferably a compression mechanism configured to
compress the ambient air. The compression mechanism is preferably a
compressor as described above, but may also be other compression
mechanisms such as aerodynamic devices that operate using the ram
pressure generated by the aircraft's airspeed. Other known oxygen sources
such as oxygen tanks are also within the scope of the invention.

[0054] With reference to FIGS. 1A, 3 and 4, the fuel tank preferably
includes an inner aluminum tank liner 151, having an inner carbon layer
153 formed on it, and an outer aluminum tank liner 155, with an outer
carbon layer 157 formed on it. The internal radius of the inner aluminum
layer is preferably four feet. Such a tank will preferably hold
approximately 1,180 pounds of liquid hydrogen.

[0055] Core cells 171 are bonded onto and extend between the inner and
outer aluminum tank liners 151 and 155 to connect them. These cells are
preferably hexagonal, having vent holes 173 in the walls of the cells. A
vacuum is created between the inner and outer aluminum tank liners,
minimizing heat transfer between the fuel and the outside environment.
The vent holes minimize the direct heat-conduction path. Preferably, each
cell extends four inches between opposing sides. The fuel tank preferably
insulates the liquid hydrogen fuel so as to receive 28 or fewer watts
through convection from the surrounding, ambient air.

[0056] The fuel cell is configured to operate at one or more
power-generation rates that require the gaseous hydrogen to be supplied
at related operating-rates of flux. The heat received by the liquid
hydrogen via convection through the insulated tank walls preferably
causes the liquid hydrogen to boil at a boiling-rate lower than one or
more (and preferably all) of the anticipated boiling-rates desired to
produce gaseous hydrogen at the related operating-rates of flux. However,
if a hybrid power system (e.g., a combination fuel cell and solar cell
system) is used, there might be times when a zero boiling rate would be
preferred.

[0057] To provide hydrogen to the fuel cell at an acceptable rate over the
convection boiling rate, heat is either delivered to, or generated in,
the fuel tank 111 by a heat source. That heat source is configured to
increase the boiling-rate of the liquid hydrogen to one or more desired
boiling-rates adequate to supply gaseous hydrogen to the fuel cell at an
operating-rate of flux. The fuel tank is configured to supply hydrogen to
the fuel cell at a rate related to and/or determined by the boiling-rate
of the hydrogen, and thus operate the fuel cell at a power-generation
rate adequate to power generation needs.

[0058] Preferably the heat source is an electrical heating element. The
fuel in the fuel tank is preferably boiled off over ten or more days to
maintain the aircraft's flight for at least that length of time.
Preferably 1.5 kilowatts of heater power are required to boil the liquid
hydrogen off over that period of time. The heater is preferably
configured such that increased levels of heater power are readily
available when needed.

[0059] The aircraft is preferably includes a controller configured to
regulate the reaction pressure of either of (or both of) the fuel and the
oxidizer. The controller regulates these reaction pressures in response
to the power requirements of the aircraft, increasing the given
power-generation rate by appropriately increasing the reaction pressures
of one or both of the reactants. The controller preferably regulates the
reaction pressure of the oxidizer by regulating the amount that the
compressor compresses the ambient air being fed to the fuel cell.
Similarly, the controller preferably regulates the reaction pressure of
the fuel being fed to the fuel cell by regulating the heat generated by
the heat source, and thus the boiling rate of the liquid fuel.

[0060] Under the regulation of the controller, at near sea-level
conditions (e.g., conditions during take-off at normal ground-level
airports), the compressor might operate to compress the ambient air to
reaction pressures above 1 atmosphere. However, compression will only
occur if it is needed for adequate power generation. As the aircraft
ascends to higher altitudes, such as cruise-level altitudes, the ambient
air is lower in pressure. Rather than increasing compression levels to
completely compensate for the lower ambient pressures, the controller
regulates the compression of ambient air to produce reaction pressures
below 1 atmosphere.

[0061] Under the regulation of the controller, at near sea-level
conditions such as take-off conditions, the heat source will generally
operate to produce hydrogen gas at a reaction pressure that might exceed
1 atmosphere. However, to maintain the fuel supply, hydrogen will only be
provided at a rate needed to generate the power required by the aircraft
from the fuel cell.

[0062] During the flight of the aircraft, the controller preferably
regulates the reaction pressures of the oxygen and the hydrogen such that
hydrogen is reacted at the minimum rate necessary for the fuel cell to
meet its power-generation rate requirements. This minimizes both fuel use
and the power drain from boiling the liquid hydrogen. The controller also
preferably regulates compressor operation to minimize power usage, which
typically means that the air is compressed as little as possible.
However, to protect the integrity of the fuel cell membrane, the reaction
pressure of the fuel is maintained at no greater than a predetermined
increment away from (and generally above) the reaction pressure of the
oxidizer.

[0063] Preferably, at stratospheric conditions (e.g., 55,000-70,000 ft),
where ambient air will typically be at less than 2 psia, the controller
will typically regulate the compressor to produce an air reaction
pressure of approximately 6 psia (i.e., 6 psia, within a range
established by significant digits). Preferably the increment of the fuel
reaction pressure from the oxidizer reaction pressure will be no more
than approximately 4 or 5 psi. Thus, at stratospheric conditions the
controller will typically regulate the fuel source to produce a fuel
reaction pressure of no more than approximately 10 or 11 psia.

[0064] Based on the recited fuel and propulsion system, it is estimated
that the aircraft, with a gross weight of 4,000 pounds, can loiter at
60,000 feet MSL within an area of 3,600 feet, with a speed of 130 feet
per second, and a potential dash speed of 180 feet per second when
necessary. To maintain a presence within the loiter diameter, the
aircraft will bank up to 15 degrees in turning maneuvers.

First Preferred Aircraft Embodiment

Airfoil Camber

[0065] With reference to FIG. 5, the wing of the preferred embodiment
preferably includes a highly cambered airfoil 201 that provides for high
lift in low-speed flight regimes. The airfoil's camber permits the
airfoil to achieve a lift coefficient of about 1.5 at the Reynolds number
typically experienced by sailplanes and the stratospheric aircraft of the
invention.

[0066] An important aspect of the use of highly cambered airfoils is that
they cause large negative pitching moments on the wing. In an alternate
variation, this embodiment uses flaps 203 to produce a highly cambered
airfoil in low-speed flight regimes (see, FIG. 6).

First Preferred Aircraft Embodiment

Variation One--Stiff Wing

[0067] In a first variation of the first preferred embodiment, the
empennage provides moments to react the overall moment of the airfoil's
negative pitching moment. The supports 118 and the wing structure in the
area outboard of the wing's connection to the slats provide structural
support and rigidity to the wing so as to avoid excessive wing torsion

First Preferred Aircraft Embodiment

Variation Two--Counteracting Moments

[0068] In a second variation of the first preferred embodiment, the wing
includes slats and/or a reflexed trailing edge, providing positive
pitching moments to react the overall effect of the airfoil's negative
pitching moment.

[0069] In this variation, the wing of the preferred embodiment includes
leading edge slats 205 that extend conventionally. The slats increase the
CL before stall, and can optionally be deployed autonomously by
relative wind directions and/or pressures. The deployment of some slat
variations can produce slight pitching moment change. While this effect
is not relevant for ordinary aircraft, it is important in the present
embodiment in preventing significant pitch-down wing torsion, such as
caused by the use of highly cambered airfoils (or flaps, as shown in FIG.
6). Furthermore, the slats let the airfoil of the invention achieve a
higher CL in slower flight, and can retract at higher speeds to cut
drag and limit gust loads.

[0070] The use of a slat 205, preferably extending autonomously in high
CL flight regimes and retracting in low CL flight regimes, can
increase the max CL as much as 0.4 or even more, while having either
a negligible or even a preferable effect on the pitching moment. With
careful design and execution, the drag at both lower and higher speeds
can be minimized, as demonstrated by airliners incorporating slat
technology.

[0071] Additionally, in this variation the wing's airfoil 201 incorporates
a reflexed portion 207 at its trailing edge, and is preferably configured
to produce a net zero or slightly positive pitching moment even though
the wing has high camber. In effect this simulates a standard
downward-loaded tail, with a very short moment arm, in the airfoil
itself. Such airfoils can achieve a high maximum CL for lower
speeds, with reasonably low drag at higher speeds, while avoiding the
wing torsion problems caused by flaps.

Other Embodiments in General

[0072] Preferred embodiments of the invention have a variety of potential
uses, the primary one being to carry a radio relay station that
facilitates communications between ground, air, and/or satellite
entities. For radio relay purposes, such embodiments must support an
antenna platform that is horizontally (and azimuthally) stabilized and
can "see" out in all directions 25° below the horizontal without a
wing or tail obstructing the view. Optionally, the antenna platform can
be lowered for use during flight and raised to avoid contacting the
ground during landings, takeoffs and taxiing.

[0073] A common role for embodiments of the aircraft will be to substitute
for solar-powered aircraft, such as the one disclosed in U.S. Pat. No.
5,810,284 (the '284 patent), that cannot stationkeep for part or all of
the year in some locations due to strong winds and/or limited solar
radiation, such as is associated with long nights and low angles of
available sunlight during the winter at high latitudes.

[0074] The preferred aircraft is unmanned, and can stationkeep closely
within a limited boundary. Being unmanned, the aircraft is preferably
controlled either by an autonomous system or by remote piloting.

[0075] In order to stationkeep closely, the aircraft will be slow flying
when winds are light and it will generally maneuver continuously. The
aircraft also will fly sufficiently fast enough to stationkeep in strong
winds, and to fly significant distances to landing fields having benign
weather conditions. It also will have enough climbing ability at peak
altitudes when fully loaded (i.e., at the early, fully fueled stage of
the flight) to maintain its altitude in atmospheric down-currents.

[0076] The preferred aircraft will ordinarily stationkeep in the vicinity
of an altitude between 55,000-70,000 feet. The available speed range will
range from a stall speed at less than 20 mph IAS (indicated airspeed) to
more than 40 mph IAS, which is about 70-140 mph TAS (true airspeed) at
65,000 feet.

Other Embodiments in General

Fuel and Power Systems

[0077] Preferred embodiments of the invention are fueled by liquid
hydrogen reacted with atmospheric oxygen in a fuel cell. This fuel
provides a high energy content. Thus, these embodiments preferably can
operate even in zero sun conditions.

[0078] Other embodiments, including variations of those described above
and below, can be configured to use other fuels, and preferably to use
gaseous fuels that are stored in liquid form.

[0079] Optionally, embodiments of the aircraft can include solar cells to
prolong its flight in conditions having extensive available solar
radiation. Furthermore, other hybrid combinations of power sources can be
used, including ones using regenerative fuel cells and/or conventionally
combusted fuels (e.g., turbines or reciprocating engines) and they are
within the scope of the invention. A conventionally combusted fuel would
preferably draw oxygen for combustion from the surrounding air (usually
with some compression).

[0080] The mechanical power generated by the power sources can directly
drive either a propeller or an attached generator that provides
electricity for propeller-driving electric motors. A generator may well
be needed for communication, control, and payload operation. A
multiple-motor control logic unit can mix power from multiple power
sources as each situation requires. Additionally, embodiments will
preferably have a small battery energy system to provide redundant power
for vehicle communication and control. This battery power can also be
used to make landing maneuvers safer.

Other Embodiments in General

Configuration

[0081] In part, the invention pertains to the overall vehicle geometry.
The configuration of each embodiment is subject to numerous tradeoff
considerations. The low-speed flight capability is preferably
accommodated through the use of low aircraft weight, large wing area and
high maximum lift coefficients of the wing airfoil. The power required at
lower speeds is minimized by using a large wingspan that reduces induced
drag. High-speed flight is preferably accommodated through the use of
higher power generation rates, lower lift coefficients of the wing
airfoil, smaller wing area, an extremely clean design and exterior
structure, as well as appropriately designed propeller(s). The shifting
of the aircraft's CG (center of gravity) and the varying of the
aircraft's rotational inertia as the fuel is consumed can be limited by
appropriate fuel tank management.

[0082] Through the use of larger airfoil chords for a given span, larger
wing areas and reduced stall speeds can be achieved. There is also a
slightly decreased power requirement, although the added weight of a
"fat" wing may negate these benefits. Nevertheless, for the preferred
role of embodiments of the present invention, a slower flight speed, even
at the cost of extra power, can decrease the extent of the maneuvering
necessary to stationkeep during low wind speeds, and thereby increase
efficiency. Depending on the operational requirements, a normal
optimization study can determine the most useful chord compromise for a
final design.

[0083] To accommodate these conflicting design criteria, and thereby
provide for a large speed range in preferred embodiments, the aircraft is
preferably characterized by a geometry change between the low-speed and
high-speed flight regimes. The extent of the speed range will vary
depending on the stationkeeping requirements. Less-stringent
stationkeeping requirements (both laterally and vertically) can permit
the aircraft to operate with more efficient, gentle turns and to move to
different altitudes if the wind profile showed a benefit in doing so,
thus requiring less of a speed range than tighter stationkeeping
requirements.

[0084] Typically, a conventional vehicle is given pitch and yaw stability
primarily by a large tail moment (the tail forces times the moment arm
between the wing and the tail) and/or by a canard in front of the wing
that, for pitch stability, operates at a higher lift coefficient than the
wing and stalls earlier. Tails mounted where they are in the up-flow of
wing tip vortices can be much smaller than normal tails positioned in the
wing downwash, but there are structural difficulties with such "outboard
tails."

[0085] As a fuel load is consumed, the aircraft's CG (center of gravity)
and rotational inertia will vary. This effect can be limited by
appropriate fuel tank management.

Other Embodiments in General

Airframe Components

[0086] In part, the invention pertains to the specific design of
aircraft's airfoils. A torsionally flexible wing is characteristic of
many embodiments of the present invention. The typical airfoil of the
invention has enough camber to permit it to achieve a lift coefficient of
about 1.5 at the Reynolds number typically experienced by the aircraft.
As noted above, there preferably is some geometry change of the aircraft
between the low-speed and high-speed flight regimes. Camber-changing
devices are relatively simple and useful devices for changing airfoil
geometry.

[0087] An important aspect of the use of either flaps or highly cambered
airfoils designed for high lift, is that such flaps (when extended
downward) and airfoils cause a large negative pitching moment on a wing.
This affects both the aircraft's overall stability and the wing's
torsional deflection. Such wing twist at the outer portions of the wing,
due to a negative pitching moment, can be a severe problem with
torsionally flexible, long-span wings such as are common in preferred
embodiments of the invention.

[0088] Regarding the aircraft's overall stability, this problem can be
handled by canard or tandem or tailed aircraft approaches within the
scope of the invention. The configuration can produce enough pitch
stability to overcome the negative pitch effect of the airfoil. The front
surface needs to have less percentage lift increase due to a small upward
gust than does the rear surface. This is accomplished by having the front
surface operate at a higher CL than does the rear. Note that the
rear surface is operating in the downwash wake of the front surface. For
the standard configuration this merely decreases the stabilizing effect
of the tail, but the vehicle is still stable. For canard configurations
the downwash effect becomes much more troublesome, and dictates much
higher CLs for the front surface than for the rear, creating both
overall vehicle inefficiencies and stall problems.

[0089] The larger problem caused by negative pitching moments is that, for
a torsionally flexible wing the wing can twist significantly under the
pitching moment. This twisting can even produce net negative lift in the
outer wing, which is the cause of the undesirable aileron-reversal
effect.

[0090] Many embodiments of the present invention incorporate flexible wing
design aspects such as those disclosed in the '284 patent. Various of
these embodiments use one or more mechanisms to counteract this problem.

[0091] As noted in the first preferred embodiment, slats provide a
mechanism to counteract negative pitching moments in some embodiments, as
well as increasing the CL (coefficient of lift) by 0.3, or even as
much as 0.4 or more before the onset of stall. Likewise, as used in the
first preferred embodiment, a reflexed airfoil further counteracts the
negative pitching moment in some embodiments. With careful design, one or
both of these mechanisms can be used to achieve a high maximum CL
for lower speeds with reasonably low drag at higher speeds. Vortex
generators can be used on the rear, underside of slats to induce vortices
that may permit still higher maximum CLs.

[0092] Other mechanisms are provided in some embodiments to limit the
effects of a negative pitching moment, as discussed in more detail in the
additional preferred embodiments below. These include "section" tails or
canards and swept flying wings.

[0093] The slatted, highly cambered and reflexed airfoil can be used in
both standard-aircraft type embodiments of the invention, such as the
first preferred embodiment, and also in flying wings. If the wing of a
flying wing is swept, it causes more pitch damping and stability. This
also makes CG changes from fuel withdrawal from elongated-fore-aft tanks
more tolerable.

Additional Preferred Embodiments

[0094] Both the first preferred embodiment and the additional preferred
embodiments below are to be understood as including variations
incorporating different combinations of the power system and aircraft
component features described in this specification. Individual details
such as the number and placement of the motors are not depicted in some
of the figures for simplicity.

Second Preferred Aircraft Embodiment

[0095] The second preferred embodiment of the invention incorporates
various combinations of the above-described features into an aircraft
incorporating the structural features of the span-loaded flying wing
disclosed and/or depicted in the '284 patent. Of particular note,
variations of this embodiment incorporate a fuel cell operating at
pressures described above with regard to the first preferred embodiment.
Additionally, variations of this embodiment incorporate a fuel cell
storage tank configured to contain liquid hydrogen, and a heater to boil
the liquid hydrogen at a determined or predetermined boiling rate.

[0096] This aircraft is characterized by very flexible wing segments that
typically have a very slight positive pitching moment by virtue of the
airfoils selected. While variations of the second preferred embodiment
can include highly cambered airfoils, flaps, slats and/or reflexed
trailing edges, this embodiment has not been found to be a highly
efficient platform for using high camber.

Third Preferred Aircraft Embodiment

[0097] With reference to FIGS. 7A and 7B, in this embodiment a wing 301 is
divided into a number of subsections 303, six being shown in the figure.
Each subsection has a tail that permits the negative pitching moments of
that section's highly cambered airfoil (or flap) to be reacted. The four
outboard sections preferably have separate tails 305, and the two inboard
sections share a laterally extending tail 307. Optionally, the sectional
structure of this preferred embodiment can adopt many of the features and
characteristics of the previous preferred embodiment and/or the aircraft
disclosed in the '284 patent.

[0098] In this multi-tail assembly, each of two symmetrically located
"bodies" or fins 309 holds a liquid hydrogen storage and fuel cell
system. Two systems are preferably used for both symmetry and
reliability. The two fins support the shared laterally extending tail
307. The two fins also support landing gear, and a communications
platform 311, which extends downward for better unobstructed viewing, can
be retracted upward for landing.

[0099] It should be noted that, as this embodiment rolls in flight, the
outboard subsections will tend to orient relative to the local flow
regime and thus decrease the roll damping. Active control of the tails on
the end subsection units can be used to eliminate the problem. However,
the use of active control systems does increase the complexity of the
system and thereby reduce its reliability.

[0100] The aircraft preferably has enough tails distributed across the
wing 301 to handle the pitching moment for each of the wing's subsections
303, providing for both vehicle pitch stability and limited wing twist.
If this embodiment's wing is designed torsionally stiff enough to keep
the wing from significantly twisting under section pitching moment
influences, then some or all of the four outer, separate tails 305 can be
removed and the central, laterally extending tail 307 can provide vehicle
pitch stability, even with flap deployment.

Fourth Preferred Aircraft Embodiment

[0101] With reference to FIG. 8, in this embodiment a conventional
aircraft layout is provided with a flexible wing 401, which supports a
fuselage 403 and is divided into a number of subsections 405. Similar to
the third preferred embodiment, each subsection has a wing-tail 407 that
permits the negative pitching moments of that section's highly cambered
airfoil (or flap) to be reacted. The main concern of the wing-tails is to
prevent local wing torsion, as the overall aircraft pitching moments can
be reacted by a tail (not shown) mounted on the fuselage.

[0102] Because the wing is flexible, the roll damping is decreased as the
wing twists during roll. This effect can be decreased if the sections
rotate on a strong spar, and both tip sections are rigidly attached to
the torsionally stiff spar to provide roll damping. The wing could be
swept in variations of this embodiment.

Fifth Preferred Aircraft Embodiment

[0103] With reference to FIG. 9, in this embodiment a long and typically
flexible wing 501, such as might be found in the third preferred
embodiment, is connected to a laterally extending tail 503 by a plurality
of small "fuselages," 505 some of which could simply be spars. The tail
extends laterally across substantially the entire wing. Two primary
fuselages 507 preferably include fuel and power modules. The aircraft
thus has pitch stability all across the span, even with the use of flaps
on the wing.

[0104] The torsional flexibility of the wing and tail sections of this
embodiment will need to be made adequately rigid enough to limit
deflection during roll unless active control is to be used. As noted
above, it is preferable to avoid active control if possible.

[0105] As previously noted for all embodiments, this embodiment can
include variations having different combinations of slats and flaps
(e.g., slotted flaps). These include variations characterized by the tail
having a small chord and zero lift at intermediate speeds.

Sixth Preferred Aircraft Embodiment

[0106] With reference to FIGS. 10A and 10B, in this embodiment a long and
typically flexible wing 601, such as might be found in the third
preferred embodiment, is connected to a laterally extending tail 603 by a
plurality (namely four) of "fuselages," 605, each being a fuel/power
module that also provides an adequate moment arm to support the tail.
Each outboard end 607 of the wing extends roughly 25 feet beyond the
outermost fuselage and is made torsionally strong enough such that
flap/aileron deflection is limited to about half of that used in the
inner, span-loaded 90 feet of the wing. This construction provides that
aileron reversal will only occur at speeds significantly higher than
preferred indicated airspeeds.

[0107] Additionally, the four fuselages 605 provide mountings for simple
landing gear (e.g., two tiny retractable wheels on each fuselage). A
radio relay pod can be lowered during flight to a level where 30°
banks of the aircraft will not obstruct the pod's visibility at more than
20° below the horizon.

Variations of the Third through Sixth Embodiments

[0108] Another approach within the scope of the invention is to vary the
above-described third through sixth embodiments to have canards rather
than tails. It should be noted that a lower CL is required on the
rear wing surface (i.e., it has an early stalling front surface). This
will likely cause higher levels of drag than the described variations
with tails.

Seventh Preferred Aircraft Embodiment

[0109] With reference to FIGS. 11A and 11B, a seventh, preferred
embodiment is a swept flying wing design having a wing 701 and a
6-element tetrahedron frame 703 formed of compression struts. A fuel and
power module 705 and a radio platform 707 are centrally located and
preferably supported by the tetrahedron frame. The tetrahedron frame adds
great strength to the inner portions of the wing, permitting the weight
of fuel and power module and the radio platform to be handled readily.
Stabilizers and/or control surfaces can optionally be mounted on the fuel
and power module to add further stability and/or control.

[0110] Three elements of the tetrahedron frame 703 are preferably in a
plane defined by the wing's main spars, extending along both sides of the
wing 701. Two wing-based elements 721 of these three spar-plane elements
either extend from a common, structurally reinforced point at the nose,
along the spars, or are composed of the spars themselves. The third of
the three spar-plane elements is a laterally extending element 723 that
extends between the spars from spar locations roughly 50 or 60 feet
apart. There can be a benefit in integrating the lateral element of the
tetrahedron into an extended wing chord in the middle portion of the
aircraft (not shown). If this embodiment's span is 140 feet (having an
aspect ratio approximately in the range of 14-17.5), the cantilevered
wing elements outboard of the tetrahedron will laterally extend 40 or 45
feet each, being a somewhat longer distance when considering the sweep,
but still a relatively short distance that is consistent with good
torsional and bending strength.

[0111] The remaining three elements of the tetrahedron frame 703 extend
downward to a common point 725. Two side-descending elements 727 of these
three downward-extending elements extend down from the two ends of the
laterally extending element 723, while the third, a center-descending
element 729, of these three downward-extending elements extends down from
the common, structurally stiff or reinforced part of the spars at the
nose of the aircraft, where the two wing-based elements 721 meet.

[0112] The drag of externally exposed compression struts is of aerodynamic
relevance, and these should be design aerodynamically. Omitting the
portions of the compression struts that are within the wing, the
remaining, exposed elements represent roughly 100 feet or less of exposed
strut length. With 1 foot chord, and a low drag shape giving a Cdo
of approximately 0.01, only 1 ft2 of equivalent flat plate area is
added to the plane by the exposed elements.

[0113] In the relatively simple configuration of this embodiment, pitch
and yaw control can be achieved by tip elevons, or more preferably, by
wingtips that rotate about an axis along the wing's quarter chord. This
rotating-tip type of control has been successfully implemented in flying
wings and conventional aircraft.

[0114] A benefit of many variations of the swept flying wing is that, by
appropriate wing twist (and hence lift distribution) the tips can be in a
region featuring upwash, letting the tips produce thrust and permitting
banked turns without causing adverse yaw. This is accomplished without
the drag of a vertical surface.

[0115] Furthermore, many variations of this embodiment will have strong
pitch stability, thus providing the ability to accommodate a reasonable
negative pitching moment, such as from positive flaps that increase
camber. If these portions are forward of the CG, the configuration pitch
stability is more readily able to accommodate the effects of airfoil
pitch instability. Preferably this embodiment of the aircraft includes a
cambered airfoil with reflex and slats, taking full advantage of the
strong tetrahedron structure for distributing loads. Thus, the
combination of the cambered/reflexed/slatted airfoil, used on the flying
wing of the present embodiment, is especially preferred

[0116] With reference to FIG. 11C, in a variation of the seventh,
preferred embodiment, two power pods 751 are located far out at the ends
of the laterally extending tetrahedron element 723, making the aircraft
into a span-loaded, swept flying wing.

A Further Variation of the Embodiments

[0117] The above described embodiments can each be varied so as to be
directed to a multi-wing aircraft such as a biplane, such as with each
wing having half the chord of the equivalent monoplane wing. The vehicle
performance would remain about the same but the wing's negative pitching
moment effect would be reduced because the chord would be halved. The big
box truss has merit for achieving torsional and bending rigidity and
perhaps for lower wing weight. Nevertheless, there is a drag penalty due
to the struts and wires and their intersections with the wing.

[0118] If a 100-foot span wing with an 8' chord and thus a 12.5 aspect
ratio (800 ft2, at a high-speed CL of 0.3 having a parasite
drag coefficient of 0.007 and hence a drag area of 5.6 ft2) were
equated to a biplane with two 4-foot chord wings, having 600 ft of 1/16''
piano wire to stabilize the box formed by the two wings, the wire drag
area would be more than 3 ft2. Considering strut drag, and the fact
that the lower Reynolds number for the airfoils adds to their drag, the
wing drag area would more than double and inhibit high-speed flight for
that embodiment.

[0119] From the foregoing description, it will be appreciated that the
present invention provides a number of embodiments of a lightweight
aircraft capable of both stationkeeping and flight over a wide range of
speeds, while consuming low levels of power, for an extended period of
time, while supporting an unobstructed communications platform, and while
exhibiting simplicity and reliability

[0120] While a particular form of the invention has been illustrated and
described, it will be apparent that additional variations and
modifications can be made without departing from the spirit and scope of
the invention. Thus, although the invention has been described in detail
with reference only to the preferred embodiments, those having ordinary
skill in the art will appreciate that various modifications can be made
without departing from the invention. Accordingly, the invention is not
intended to be limited, and is defined with reference to the following
claims.